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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Longitudinal aerodynamic performance of a series of power-law and minimum wave drag bodies at Mach 6 and several Reynolds numbers

Experimental data have been obtained for two series of bodies at Mach 6 and Reynolds numbers, based on model length, from 1.4 million to 9.5 million. One series consisted of axisymmetric power-law bodies geometrically constrained for constant length and base diameter with values of the exponent n of 0.25, 0.5, 0.6, 0.667, 0.75, and 1.0. The other series consisted of positively and negatively cambered bodies of polygonal cross section, each having a constant longitudinal area distribution conforming to that required for minimizing zero-lift wave drag at hypersonic speeds under the geometric constraints of given length and volume. At the highest Reynolds number, the power-law body for minimum drag is blunter (exponent n lower) than predicted by inviscid theory (n approximately 0.6 instead of n = 0.667); however, the peak value of lift-drag ratio occurs at n = 0.667. Viscous effects were present on the bodies of polygonal cross section but were less pronounced than those on the power-law bodies. The trapezoidal bodies with maximum width at the bottom were found to have the highest maximum lift-drag ratio and the lowest mimimum drag.

Ashby, G. C., Jr.↗
NASA NTRS1986

Transonic airfoil analysis and design in nonuniform flow

A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.

Chang, J. F.↗
NASA NTRS1960

A Method of Reducing Aeroelastic Effects of Highly Swept Wings

The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have < been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.

Hopkins, Edward J.↗
NASA NTRS1952

Pressure distribution and aerodynamic coefficients associated with heat addition to supersonic air stream adjacent to two-dimensional supersonic wing

The modifications in the pressure distributions and the aerodynamic coefficients associated with additions of heat to the two-dimensional supersonic in viscid flow field adjacetnt to the lower surface of of a 5-percent-thickness symmetrical circular-arc wing are presented in this report. The pressure distributions are obtained by the use of graphical method which gives the two-dimensional supersonic inviscid flow field obtained with moderate heat addition. The variation is given of the lift-drag ratio and of the aerodynamic coefficients of lift, drag, and moment with free stream Mach number, angle of attack, and parameters defining extent and amount of heat addition. The six graphical solutions used in this study included Mach numbers of 3.0 and 5.0 and angles of attack of 0 degrees and 2 degrees.

Pinkel, I Irving↗
NASA NTRS1960

Motion and Heating During Atmosphere Reentry of Space Vehicles

The results of an analysis of the motion and heating during atmospheric reentry of manned space vehicles has shown the following: 1. Flight-corridor depths which allow reentry in a single pass decrease rapidly as the reentry speed increases if the maximum deceleration is limited to 10 g. 2. Use of aerodynamic lift can result in a three-to five fold increase in corridor depth over that available to a ballistic vehicle for the same deceleration limits. 3. Use of aerodynamic lift to widen these reentry corridors causes a heating penalty which becomes severe for values of the lift-drag ratio greater than unity for constant lift-drag entry. 4. In the region of most intense convective heating the inviscid flow is generally in chemical equilibrium but the boundary-layer flows are out of equilibrium. Heating rates for the nonequilibrium boundary layer are generally lower than for the corresponding equilibrium case. 5. Radiative heating from the hot gas trapped between the shock wave and the body stagnation region may be as severe as the convective heating and unfortunately occurs at approximately the same time in the flight.

Wong, Thomas J.↗
NASA NTRS1960

Static Longitudinal Stability and Control Characteristics At A Mach Number of 1.99 of a Lenticular-Shaped Reentry Vehicle

An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.99 to determine the longitudinal stability and control characteristics of a reentry model consisting of a lenticular-shaped body with two fin configurations (horizontal fins with end plates). Effects of deflecting the larger size fins as pitch-control surfaces were also investigated. The results indicate that the body alone was unstable from an angle of attack of 0 deg to about 55 deg where it became stable and remained so to 90 deg. The addition of fins provided positive longitudinal stability throughout the angle-of-attack range and increased the lift-drag ratio of the configuration. Reducing the horizontal-fin area at the inboard trailing edge of the fin had only a small effect on the aerodynamic characteristics of the vehicle for the condition of no fin deflection. Deflecting the fins, appeared to be an effective means of pitch control and had only a small effect on lift-drag ratio.

Jackson, Charles M., Jr.↗
NASA NTRS1977

The oblique wing - Aircraft design for transonic and low supersonic speeds

In tests at transonic speeds a conventional 'subsonic' wing turned at an oblique angle to the flow has demonstrated lift-drag ratios higher than conventional swept-back or delta wings. When returned to its straight position, perpendicular to the flow, such a wing would minimize the display of energy and noise in the airport environment. Following a brief review of theoretical and experimental results, the paper discusses various characteristics of the oblique wing as they relate to aircraft design. Topics covered are lift-drag ratio, flight control and trim and aeroelastic stability. The possibility of an all-wing aircraft for flight at Mach 2 is suggested.

Jones, R. T.↗
NASA NTRS1961

Analysis of X-15 Landing Approach and Flare Characteristics Determined from the First 30 Flights

The approach and flare maneuvers for the first 30 flights of the X-15 airplane and the various control problems encountered are discussed. The results afford a relatively good cross section of landing conditions that might be experienced with future glide vehicles having low lift-drag ratios. Flight-derived drag data show that preflight predictions based on wind-tunnel tests were, in general, somewhat higher than the values measured in flight. Depending on configuration, the peak lift-drag ratios from flight varied from 3.5 to 4.5 as compared with a predicted range of from 3.0 to 4.2. By employing overhead, spiral-type patterns beginning at altitudes as high as 40,000 feet, the pilots were consistently able to touch down within about +/-1,000 feet of a designated point. A typical flare was initiated at a "comfortable" altitude of about 800 feet and an indicated airspeed of approximately 300 knots., which allowed a margin of excess speed. The flap and gear were extended when the flare was essentially completed, and an average touchdown was accomplished at a speed of about 185 knots indicated airspeed, an angle of attack of about 7 deg, and a rate of descent of about 4 feet per second. In general, the approach and landing characteristics were predicted with good accuracy in extensive preflight simulations. F-104 airplanes which simulated the X-15 landing characteristics were particularly valuable for pilot training.

Matranga, Gene J.↗
NASA NTRS1972

Effect of configuration modification on the hypersonic aerodynamic characteristics of a blended delta wing-body entry vehicle

The longitudinal, lateral, and directional aerodynamic characteristics of a delta-wing configuration were obtained experimentally at Mach 20 in helium with Reynolds numbers, based on model length, of 1.5 million and 2.9 million and at a Mach number of 6 in air with a Reynolds number, based on model length, of 4.8 million. The angles of attack varied from 0 deg to 55 deg for two sideslip angles. The effects of the addition of dorsal fins, the removal of wing tip fins, an increase in elevon span, and changes in elevon hinge-line sweep angle are discussed. The unmodified vehicle had a maximum lift-drag ratio of 2.1 at Mach 19 and of 2.4 at Mach 6 with about the same lateral and directional stability level at both Mach numbers. As the Mach number increased from 6 to 20, the longitudinal center of pressure moved forward and more positive elevon deflection was therefore required to maintain a given trim angle. The removal of wing tip fins increased the maximum lift-drag ratio and had a negligible effect on longitudinal stability, but caused directional instability that was not corrected by the dorsal fins examined. The shape of the wing and elevon hinge-line sweep had a large influence on the induced yawing moment due to roll control.

Arrington, J. P.↗
NASA NTRS1961

Flight-Path Characteristics for Decelerating From Supercircular Speed

Characteristics of the following six flight paths for decelerating from a supercircular speed are developed in closed form: constant angle of attack, constant net acceleration, constant altitude" constant free-stream Reynolds number, and "modulated roll." The vehicles were required to remain in or near the atmosphere, and to stay within the aerodynamic capabilities of a vehicle with a maximum lift-drag ratio of 1.0 and within a maximum net acceleration G of 10 g's. The local Reynolds number for all the flight paths for a vehicle with a gross weight of 10,000 pounds and a 600 swept wing was found to be about 0.7 x 10(exp 6). With the assumption of a laminar boundary layer, the heating of the vehicle is studied as a function of type of flight path, initial G load, and initial velocity. The following heating parameters were considered: the distribution of the heating rate over the vehicle, the distribution of the heat per square foot over the vehicle, and the total heat input to the vehicle. The constant G load path at limiting G was found to give the lowest total heat input for a given initial velocity. For a vehicle with a maximum lift-drag ratio of 1.0 and a flight path with a maximum G of 10 g's, entry velocities of twice circular appear thermo- dynamically feasible, and entries at velocities of 2.8 times circular are aerodynamically possible. The predominant heating (about 85 percent) occurs at the leading edge of the vehicle. The total ablated weight for a 10,000-pound-gross-weight vehicle decelerating from an initial velocity of twice circular velocity is estimated to be 5 percent of gross weight. Modifying the constant G load flight path by a constant-angle-of-attack segment through a flight- to circular-velocity ratio of 1.0 gives essentially a "point landing" capability but also results in an increased total heat input to the vehicle.

Luidens, Roger W.↗
NASA NTRS1961

Experimental Investigation of a Hypersonic Glider Configuration at a Mach Number of 6 and at Full-Scale Reynolds Numbers

The aerodynamic characteristics of a hypersonic glider configuration, consisting of a slender ogive cylinder with three highly swept wings, spaced 120 apart, with the wing chord equal to the body length, were investigated experimentally at a Mach number of 6 and at Reynolds numbers from 6 to 16 million. The objectives were to evaluate the theoretical procedures which had been used to estimate the performance of the glider, and also to evaluate the characteristics of the glider itself. A principal question concerned the viscous drag at full-scale Reynolds number, there being a large difference between the total drags for laminar and turbulent boundary layers. It was found that the procedures which had been applied for estimating minimum drag, drag due to lift, lift curve slope, and center of pressure were generally accurate within 10 percent. An important exception was the non-linear contribution to the lift coefficient which had been represented by a Newtonian term. Experimentally, the lift curve was nearly linear within the angle-of-attack range up to 10 deg. This error affected the estimated lift-drag ratio. The minimum drag measurements indicated that substantial amounts of turbulent boundary layer were present on all models tested, over a range of surface roughness from 5 microinches maximum to 200 microinches maximum. In fact, the minimum drag coefficients were nearly independent of the surface smoothness and fell between the estimated values for turbulent and laminar boundary layers, but closer to the turbulent value. At the highest test Reynolds numbers and at large angles of attack, there was some indication that the skin friction of the rough models was being increased by the surface roughness. At full-scale Reynolds number, the maximum lift-drag ratio with a leading edge of practical diameter (from the standpoint of leading-edge heating) was 4.0. The configuration was statically and dynamically stable in pitch and yaw, and the center of pressure was less than 2-percent length ahead of the centroid of plan-form area.

Seiff, Alvin↗
NASA NTRS1986

An experimental study of the lift, drag and static longitudinal stability for a three lifting surface configuration

The experimental procedure and aerodynamic force and moment measurements for wind tunnel testing of the three lifting surface configuration (TLC) are described. The influence of nonelliptical lift distributions on lift, drag, and static longitudinal stability are examined; graphs of the lift coefficient versus angle of attack, the pitching moment coefficient, drag coefficient, and lift to drag ratio versus lift coefficient are provided. The TLC data are compared with the conventional tail-aft configuration and the canard-wing configuration; it is concluded that the TLC has better lift and high-lift drag characteristics, lift to drag ratio, and zero-lift moments than the other two configurations. The effects of variations in forward and tail wind incidence angles, gap, stagger, and forward wind span on the drag, lift, longitudinal stability, and zero-lift moments of the configuration are studied.

Ostowari, C.↗
NASA NTRS1960

Experimental Investigation of Two Low-Drag Supercavitating Hydrofoils at Speeds up to 200 Feet per Second

An experimental investigation has been made in the Langley highspeed hydrodynamics facility to determine the force and moment characteristics of two hydrofoils (one having an aspect ratio of 1 and the other having an aspect ratio of 3) designed to have improved lift-drag ratios when operating under either supercavitating or ventilated conditions. Measurements were made of lift, drag, and pitching moment over a range of angles of attack from 40 to 200 for depths of submersion varying from 0 to approximately 1 chord. The range of speed for the investigation was from 110 to 200 feet per second. When the upper surface of the hydrofoils was completely unwetted, the experimental values of lift and drag forces were in good agreement with the theoretical values obtained from the zero-cavitation-number theory. The theoretical values for minimum angle of attack for operation with the upper surface of the hydrofoil unwetted define the lower limits of angle of attack for which the experimental values of lift coefficient are either in agreement with or slightly greater than those predicted by theory.

Christopher, Kenneth W.↗